Gas turbine engines, whether ground based power generation turbines or aircraft engines, produce undesirable nitrogen oxides as a result of the high temperatures associated with combustion of fuel in the burner section thereof, commonly referred to as the combustor. Naturally, it is desirable to reduce the emission of nitrogen oxides. Although it is possible to catalytically reduce nitrogen oxides back to nitrogen and oxygen, it is more effective to reduce nitrogen oxide emissions by reducing the formation of nitrogen oxides during combustion.
It is well known that the formation of nitrogen oxides during combustion of fuel in the combustor of a gas turbine engine at high power may be reduced by sectioning the combustor into two or more combustion chambers disposed in series such that the products of combustion formed in the upstream chamber pass through the downstream chamber. For example, U.S. Pat. No. 4,045,956 discloses a two zone combustor wherein a first portion of the fuel is burned in a pilot zone to provide a flow of hot gases into which additional fuel is injected before additional combustion air is added to complete combustion in a second zone of the combustor.
While effective in stretching out the combustion process so as to reduce peak temperatures and consequently nitrogen oxide formation, the use of multiple combustion chambers complicates operation, particularly if fuel is to be supplied independently to each of the combustion chambers. In order to do so, it is known to provide a fuel splitter valve in the fuel supply system of the gas turbine engine, the fuel splitter valve being positioned downstream of a fuel metering unit (FMU) which meters total fuel flow to the turbine engine in response to various operating parameters, for example power level and engine speed. The fuel splitter valve receives the total metered fuel flow from the FMU and functions to split the received fuel flow into a plurality of output flows, one per each combustion chamber. Each output flow is directed to a fuel supply manifold associated with a particular combustion chamber for distribution to the individual fuel nozzles associated with that combustion chamber and fuel supply manifold.
U.S. Pat. No. 4,949,538 discloses a system for supplying gaseous fuel to a combustor of a power generation type gas turbine wherein the gaseous fuel is supplied to two distinct zones of the combustor. The system includes a fuel flow splitter assembly which receives a metered gas flow from a main flow control and splits the received gas flow into two streams, one for delivery to a primary zone of the combustor and the other to a secondary zone of the combustor of the turbine engine. The fuel flow splitter comprises a coordinated valve splitter assembly having a linear trim valve controlling the primary gas flow in parallel with an equal-percentage trim valve controlling the secondary gas flow. The proportion of the fuel flow output to the primary and secondary zones is dependent upon the relative flow area through the primary and secondary valves, respectively. In one embodiment, coordination of the primary and secondary valves is accomplished through rigid mechanical connection of the valve shafts. Coordination may also be accomplished through shaped cams actuating the primary and secondary valves or through individual hydraulic actuators associated with each of the primary and secondary valves and an electronic control effective for appropriately shaping the control characteristics of each of the hydraulic actuators. With the primary and secondary valves being coordinated, the change in fuel flow to the primary and secondary zones may be undesirably slow during a rapid transition such as may be necessary in an aircraft turbine engine.